Gas turbine engines may be used to power various types of vehicles and systems. A gas turbine engine may include, for example, five major sections: a fan section, a compressor section, a combustor section, a turbine section, and an exhaust section. The fan section includes a fan that induces air from the surrounding environment into the engine and passes a fraction of this air toward the compressor section. The compressor section raises the pressure of the air it receives from the fan section and directs the compressed air into the combustor section where it is mixed with fuel and ignited. The high-energy combustion mixture then flows into and through the turbine section, thereby causing rotationally mounted turbine blades to rotate and generate energy. The combustion mixture exiting the turbine section is exhausted from the engine through the exhaust section. The remaining fraction of air induced into the fan section is transitioned through a bypass plenum and exhausted through a mixer nozzle.
The fan includes at least one rotor. In some engines, the compressor section is implemented with a combination of one or more axial compressors, mixed flow and/or centrifugal stages. An intermediate-pressure (IP) compressor (also known as a “booster”) may be included in some engines between the fan and a high-pressure (HP) compressor, to supercharge or boost the HP compressor helping to raise the overall pressure ratio of the engine cycle to higher levels. Both the IP and HP compressors have one or more stages comprised of a rotor and a stator. Each of the fan and compressor rotors has a plurality of rotor blades. The plurality of rotor blades are spaced in a circumferential direction around a rotor hub 2 located coaxially around a longitudinal centerline axis. Each rotor blade 1 has a pressure sidewall 3 and a circumferentially opposing suction sidewall 4 extending in a radial direction between a root 6 and a tip 7 and in an axial direction between the leading edge 9 and a trailing edge 11. As the fan and compressor rotors rotate, the rotor blades add energy to the flow and increase total pressure while the downstream stator diffuses the flow and raises the static pressure. “Stall margin” is defined as the difference between operation of the compressor and the conditions that would be required to cause stall to occur. During design, enough margin is required to insure the fan and compressors will not stall anywhere in the allowable operating range. Engine surge can result from encountering stall and, if not properly addressed, may adversely impact engine performance, durability, and flight safety. During compressor operation, stall occurs when the stream momentum imparted to the air by the rotor blades is insufficient to overcome the pressure rise across the compressor, resulting in a reduction, local reversal, or instabilities in the compressor airflow. These aerodynamic flow conditions can cause the engine to surge.
Therefore, fundamental in fan and compressor design for aircraft gas turbine engines is efficiency in compressing the air with sufficient stall margin over the entire flight envelope of operation from takeoff, cruise, and landing. However, compressor efficiency and stall margin are normally inversely related with increasing efficiency corresponding with a decrease in stall margin. The conflicting requirements of stall margin and efficiency are particularly demanding in high performance aircraft gas turbine engines that require increased cycle pressure ratio and increased efficiency while maintaining adequate stall margin for safe operation. In conventional designs, efficiency is usually sacrificed in order to achieve required stall margin and required operability.
Mechanisms have been devised to extend the stable operating range of the fan and compressor (i.e., extend the range to stall). For example, rotor casing treatments are added after design of the rotor has been completed to increase the stall margin, but usually are associated with an efficiency penalty. The rotor casing treatments are conventionally added after rotor design if the stall margin is too low. Conventional “forward tip swept rotor blades” (a single forward tip swept rotor blade 1 is depicted in FIG. 1) of a conventional rotor 5 (FIG. 2) (e.g., a fan rotor or a compressor rotor) also increase overall stall margin and have aerodynamic advantages as compared to neutral tip swept rotor blades, but at an efficiency penalty. The “forward tip swept rotor blade” 1 includes an aft sweep extending over most of the rotor blade leading edge 9 with a local zone of forward sweep (encircled region A of FIGS. 1 and 2) near the tip 7 (from about 75% to 100% of the blade span). As used herein, the term “blade span” refers to the geometry of the airfoil 1 or rotor blade that is defined in part by a span dimension S extending radially from the root 6 to the tip 7. Aerodynamic “sweep” is a conventional parameter represented by a sweep angle which is a function of the direction of the incoming air and the orientation of the airfoil surface in the axial, radial, and circumferential or tangential directions. The rotor direction of rotation is identified by arrow C in FIGS. 1 and 2. A forward sweep is denoted by a negative value for the sweep angle. A neutral sweep has zero sweep angle. An aft (or rearward) sweep is denoted by a positive value for the sweep angle. Unfortunately, the transition from the aft sweep along the blade span to the forward sweep (encircled region A) near the tip in conventional forward tip swept rotor blades creates a local transition zone (encircled region B of FIG. 1) of reduced efficiency and reduced range to stall. The forward sweep near the tip is also a mechanical design challenge with an associated risk of forced response and flutter.
Accordingly, it is desirable to provide rotors with stall margin and efficiency optimization and methods for improving gas turbine engine performance therewith. Gas turbine engine stability is increased and specific fuel consumption and turbine operating temperatures in both steady state and transient operations are decreased by using such rotors in components for gas turbine engines, including those operating at high cycle pressure ratios. Furthermore, other desirable features and characteristics of the present invention will become apparent from the subsequent detailed description of the invention and the appended claims, taken in conjunction with the accompanying drawings and this background of the invention.